1 Cal Poly Pomona Rocketry NASA Student Launch Competition POST LAUNCH ASSESMENT REVIEW April 24, 2017 California State Polytechnic University, Pomona 3801 W Temple Ave, Pomona, CA Department of Aerospace Engineering
3 Table of Contents 1.0 Overview Motor Used Altitude Reached Payload Descriptions Vehicle Summary Vehicle Dimensions Data Analysis and Results of the Vehicle Primary Payload Summary Results & Analysis Scientific Value Secondary Payload Summary Data Analysis and Results of the Payload and Scientific Value Visual Data Observed Lessons Learned Parachute Lessons Recovery Avionics Lessons Overall Lessons Summary of overall experience Educational Engagement Summary Budget Summary... 15
4 1.0 Overview 1.1 Motor Used As described in our post-frr documentation submitted to the NASA review team. The Cal Poly Pomona team used a L-1420R-P motor for the official competition launch. The L-1420R-P is an 80% L reloadable composite motor from Aerotech. The weight of the propellant was 5.64 lbs. and the total weight of the motor was lbs. The motor had a peak thrust of lbs., and average thrust of lbs., and its total impulse was lb.-sec. It had a total burn time of 3.24 seconds. 1.2 Altitude Reached Ultimately this motor carried the launch vehicle and its two payloads to an altitude of 5,793 ft., overshooting the target apogee by 513 ft. At every stage of the design process, the launch vehicle was engineered to be as aerodynamic as possible in the required sub-sonic flow regime and this may have been part of the reason that our launch vehicle significantly overshot the target apogee. The nose cone was engineered to have an elliptical shape, the fins were shaped to resemble the NACA 0012 airfoil, and the launch vehicle had as few external protuberances as possible, all in order to reduce drag and increase aerodynamic performance. Ultimately the launch vehicle could have benefitted from some additional drag to reduce its apogee. 1.3 Payload Descriptions The primary payload onboard the launch vehicle was the roll induction system. This payload took the form of a control system that had control authority over the angular motion of the launch vehicle by precisely deflecting the ailerons on the fins. The data control system samples the roll rate of the launch vehicle prior to motor burn-out. After motor burn-out, the control system activates a specific deflection of the ailerons in order to induce an angular velocity. After a predetermined amount of time, the hard-coded deflections cease and the PID control system activates, stabilizing the roll. Using the sampled roll rate prior to motor burn-out as a reference the control system returns the launch vehicle to this exact passive roll rate. As will demonstrated later in the report, the roll induction system performed exactly as designed on competition launch day. The secondary payload onboard the launch vehicle took the form of a fragile material protection system. The payload was a 3D printed Pill suspended from elastic surgical tubing to help shield the fragile materials from any sudden impulses. The interior of the payload contained foam to shield the fragile materials from the sides of the 3D printed pill and from any other fragile materials inside the pill. Unfortunately, only one of the five skeet disks survived the flight on competition day. This is entirely due to the fact that the main parachute became tangled on the descent and upon hitting the ground, the launch vehicle had an enormous amount of kinetic energy. I have full faith that all five skeet disks would have survived had the main parachute deployed properly.
5 2.0 Vehicle Summary The length of the launch vehicle from the tip of the nose cone to the end of the motor bay is 8 ft., 9 in. The outer diameter of the body tube is 6.16 in. and the total mass of the launch vehicle including a loaded motor is 45.2 lbs. The launch vehicle utilizes an Aerotech L1420 motor, whose documented performance allows for reasonable estimates for the apogee, launch acceleration, and rail exit velocity. The recovery system, utilizing two altimeters, initially activates at apogee by firing the aft ejection charge to release the cruciform drogue parachute. Once the launch vehicle has been decelerated, stabilized, and has fallen to an altitude of 700 ft., the fore ejection charge will fire, releasing the toroidal main parachute, which is predicted to inflate at 500 ft. For the launch vehicle to reach a velocity of at least 50 ft/s and therefore achieve stability, a 12 ft aluminum launch rail will be used to reach 95.5 ft/s by launch rail clearance. The primary payload takes the form of a Roll Induction System (RIS). Specifically, the RIS primarily consists of an autonomous aileron system which, following motor burnout, initiates two complete rotations and a counter rotation that ceases all angular velocity instigated by the active system. The secondary payload, the Fragile Material Protection system (FMP) consists of a pill housing suspended within a payload bay whose purpose is to shield the provided fragile materials from the loads and impulses generated by the launch and recovery of the launch vehicle. Below is figure which shows the rocket in full flight configuration and Figure with dimensions. 2.1 Vehicle Dimensions Figure Full Rocket and Subsystems Figure Full Rocket and Dimensions
6 2.2 Data Analysis and Results of the Vehicle On launch day, the rocket performed just as expected with some unexpected problems. Upon ignition, the motor burned and the rocket s velocity at rail exit was like expected values allowing the rocket to fly on its course. The longest section of body tube could withstand the peak thrust of the motor and withstood against crippling. The fastening screws used for the bulkheads were also able to handle the peak thrust and no shear out or tearing was found from the fastener locations. Due to aerodynamics, the rocket experienced a very small natural roll which was also expected. After main engine cut off (MECO) the fins were able to handle the aileron deflections which allowed the rocket to turn about its roll access which completed our primary payload criteria as well as the RFP. At apogee, the rocket was at a higher altitude than previous planned for. One cause of this was due to the aerodynamic drag of the ailerons which were higher during our test launch because the ailerons were deflected at 35 degrees for 10 seconds, but at the competition they were deflected at 12 degrees for only 2 seconds. Another underlining cause could also have been the different motor we used. The simulations that we had run said the altitude was under the flight limit, but this was not the case in life. We are still unsure what had caused our simulation to predict a much lower altitude. The ejection system worked just as planned and the drogue was released at apogee, however within seconds of descend the drogue became tangled and balled up in air. This caused the vehicle to have a much higher velocity at main deployment. The main deployed also at the right time however because of the velocity the chord zippered the side of the rocket tube, also the main parachute became tangled within its self upon descent. This all lead to a hard landing of the rocket. Upon inspection of the rocket the damage was apparent and the body tube as well as some internal bulkheads, centering rings, and fins were damaged beyond repair. However much of the electronics, internal payload bays, and motor bay were minimally damaged but mostly salvageable. Discussion of this unfortunate circumstance was debated and not altogether agreed upon by everyone on the team. One issue could have been the parachute from Fruity Chutes having too many lines on them which can cause lines to tangle. Another cause could have been bad parachute packing used to pack the parachute. Another cause could have been an undesirable packing technique used. This will be discussed in later sections. It is clear however, whatever the cause to the parachute tangling was, a main underlining problem was lack of full scale flights. Although the team had planned on four full scale flights, bad weather for two months as well as flight windows being cancelled by Edwards Airforce Base had caused the ultimate one-flight that we relied all our test data on. For future teams, it would be beneficiary to conduct more flight in order to identify problems such as these before the competition.
7 3.0 Primary Payload Summary The final flight of our Roll Induction System ( RIS ) primary experiment on April 8 th, 2017 represented the culmination of 9 months of hard work involving design, manufacturing, testing, and debugging. We are pleased to announce that the system performed beautifully on that day, and fully achieved the objectives set forth in section 3.3 of Experiment Requirements of the NSL 2017 handbook. Simplicity and reliability were the key design philosophies behind the RIS. By utilizing a mechanically coupled dual-servo configuration, redundancy and safety were incorporated into the design which guarded against an induced pitching moment. An Arduino Nano served as the system controller, which received acceleration and gyroscopic inputs from an Adafruit 9DOF sensor at a sampling rate of > 200 Hz while actuating the servos accordingly. Acceleration and gyroscopic data were recorded with our Data Collection System, also located within the Payload Bay, at a sampling rate of ~27 Hz. Video of the entire flight was captured with our Raspberry Pi Zero based Observation System. The largest uncertainty on launch day was how well our self-coded closed loop PID function would perform. Our first full scale flight test on March 4 th established the baseline functionality of the system but did not include programming which would return the rocket to its natural roll rate. On April 8 th, the previously untested PID function performed its job flawlessly without any oscillatory or divergent behavior. 3.1 Results & Analysis
8 Using angular velocity data collected by the DCS, roll rate versus time since launch was plotted. At approximately 3.1s after launch, a counter-clockwise roll is clearly induced with 17 counter deflections held for 2.5s. A brief clockwise-induced spin was held for 0.3s to provide dampening, after which the closed loop PID function took over. The PCS then sought to maintain the target rate of 3.83 rad/s recorded at burnout. While initially successful (within 10% of the target rate), as the plot shows, the PCS becomes less effective at maintaining that rate from about s. This is due to hard-coded deflection limits (18 ) that were intended to mitigate the effects of possible oscillatory behavior. These limits, combined with the lower freestream velocity during this portion of the flight, meant that the system could no longer maintain the target roll rate. This was the trade-off made between control system behavior uncertainty and actual performance. The area underneath the induced angular velocity was summed to calculate that the system induced roll, which was found to be at least 12.6 revolutions. There is some uncertainty to the exact amount due to the sensor saturation present from s, but is of little consequence since the data clearly shows the requirement was met far and beyond the minimum. A captioned video of the Roll Induction experiment is available for viewing on CPP-NSL s YouTube account here.
9 3.2 Scientific Value Besides validation of our design and testing process, the data collected during the April 8 th launch allowed us to finally calculate a CL for our aileron fins (our low speed wind tunnel suffered a breakdown just one week prior to our scheduled test in February, while data from our first full scale test suffered from sensor saturation). Using the non-saturated steady state roll rate seen during s, an angular acceleration value was found. With this, induced torque was found, as was induced lift. Using the average velocity during that time period (465.5 ft/s), a CL of 0.84 was calculated. This value is of a similar magnitude when compared with the best-case scenario 2-D predictions found using the application JavaFoil, which are shown below.
10 4.0 Secondary Payload Summary The Fragile Material Protection System, or FMP, failed to protect four out of the five fragile objects. The 3D printed plastic shell that contained the foam the fragile objects was buried in, cracked in half and separated. This fracture in the shell lead to the object experiencing unintended accelerations and breaking. The fracture was caused by the free fall landing of the bottom third section of rocket striking into the ground FMP bay first. This direct impact into the payload caused the plastic pill to max out the planned for length of the surgical tubing and the plastic pill hit the internal wall causing the fracture in the plastic shell. The remains of the fragile material can be seen in the Figure below. Figure Fragile Material 4.1 Data Analysis and Results of the Payload and Scientific Value As mentioned in section 4.0 Secondary Payload Summary, four out of the five objects survived. This result of the experiment is less than ideal, but the landing of our rocket was fittingly less than ideal. The surgical tubing that suspended the plastic shell did not snap in the abrupt landing, only over extended. Keeping the same fragile material system and the same landing condition and then thinking of how to prevent the breaking of the fragile objects the data from the landing yields the solution of simply tying the surgical tubing tighter. Making the tubing tighter would prevent the plastic shell containing the object from striking the internal wall and this added resistance will give the fragile materials a greater chance of survival. Even though there is failure in the payload there is much that can be learned, most importantly how to make the next payload better. 5.0 Visual Data Observed Two sources of data were used on Launch day, which is the data collection system and the observation subsystem that recorded the launch. The video was recorded using the observation subsystem that consists of a Raspberry Pi Zero and a Raspberry Pi Camera V2. The video started recording along with the data collections systems startup and recording a few minutes on the pad. Then launch occurred and was recorded. The video caught all of the phases of the flight. It caught the roll induction and is a secondary verification that the roll was performed. A piece of the launch video can be seen right after launch occurred. Two other rockets can be seen awaiting their launch, Figure
11 Figure Rocket Right after Launch It can be seen in the video that the rocket has a slight roll coming out of the burn phase and the roll is performed with more than the required 2 revolutions. Then the control system brings it back to its near passive roll. While, there is some steady state error included in bringing it back to the passive roll, this is most likely due to the linearization of the controller and the low velocities that the rocket was when it was attempting to get back to the passive roll. After, the rocket reaches burnout the jolt produced by the ejection of the drogue chute can be seen. It can be seen later that the drogue chute collapses in on itself due to tangled lines and the rocket goes into any uncontrolled roll. Then a jolt from the main chute deploying can be seen, but then shortly after the video cuts out. This is most likely due to power failure in the payload control system from jolting. 6.0 Lessons Learned 6.1 Parachute Lessons The recovery system suffered from a minor failure as well as a critical failure. The drogue parachute s swivel ended up seizing around 1000 ft. agl which caused the parachute to tangle and slowly crawly up the recovery line to the bulkhead show in Figure
12 This resulted in the aft body tube section becoming the main drag inducing body. This would create a rough terminal velocity of 111 ft/s assuming the aft section was perpendicular to the wind. Despite the increased speed the main parachute still should have deployed successfully and not damaged the body tube with shearing force. Figure Drogue Parachute Failure The main parachute suffered from a critical failure due to no inflation of the main. As determined from video evidence, the main parachute s initial ejection charge failed to deploy the parachute, and the secondary at 500 ft as shown in Figure was necessary to achieve deployment. This is a potential cause for the parachute lines getting tangled preventing the parachute from inflating. The inner suspension lines crossed the outer lines causing a knot, shown in Figure 6.1-3, in yellow, at the skirt which sealed the parachute entirely. Beyond the secondary deployment charge being necessary, another likely cause for this was due to a faulty parachute packing method. Figure Main Parachute Deployment and failed inflation
13 6.2 Recovery Avionics Lessons Figure Parachute knot The Primary lesson learned from this Experience is that Lipo Batteries should not be used for powering and igniting E-matches. The primary reason Lipos should not have been used is the fact that their internal resistance is too low. Repeated ejections and E-match tests show that Lipos generated, in excess, of 8 Amps through the firing circuit, putting unnecessary wear on the altimeter. Furthermore, the E-matches have shown reliable ignition as low as 1.5 V. This means the high voltage of the Lipo was unnecessary and pushed the altimeter out of its optimally designed operation voltage of 9.0 V. After this discovery, a test with varying firing circuit resistances and a 9V showed that the 9 Volt battery supplied 2 Amps for every test, regardless of circuit resistance. Lipos also pose another risk, fire hazard. Overall, they require more care to maintain, but in case of a rocket recovery failure the chance of cells puncturing increases. Upon crash the lipo could short and vent hydrogen simultaneously, resulting in the possible loss of the rocket due to fire or start a fire in an arid fuel ridden region. The second lesson learned was ease of electronics bay preparation. The switches were mounted to the body of the rocket, which required the altimeters to be disconnected to remove the sled. To mitigate this the switches could be mounted to the sled so the whole unit could be cleanly removed and reduce wiring. Another problem with the electronics bay was the bulk head design. The bulk head fit recessed completely within the coupler tube, when ejection was fired corrosive gasses would blow by the bulk head and into the electronics compartment. The solution was to cover the gap with a tacky adhesive. I propose the bulkheads should be stepped in future designs as shown in Figure Figure Stepped electronics bay Bulkhead
14 6.3 Overall Lessons One of the major lessons that was learned this year (and last year) is that retention of individuals and experience on the team from year to year is extremely important. Only one member of last year s team carried over to the team this year. This caused difficulties for the rest of the team members with regard to getting oriented to the specifics of the competition. Additionally, very few of the team members had experience with the design and manufacture of high powered rockets, which created a significant learning curve during the preliminary and critical design phases of the project. For future reference, it is very important for a larger number of team members to stay with the team from one year to the next in order to reduce these difficulties. Another important lesson that was learned this year is to focus slightly less on designing and building a launch vehicle completely from scratch and to focus more time and energy into creating original and innovative scientific payloads. One of the most challenging aspects of the entire competition is the design of payloads and it is important that the amount of time spent working on them reflects how critical they are. The payloads that the team designed and manufactured this year were incredible, but that doesn't mean they wouldn't have benefitted from more time and effort being spent on them instead of the launch vehicle. High-Powered rockets have been done many times before but innovative and original payloads are always novel. The subsystem that ultimately crippled the launch vehicle on launch day was the recovery system, which indicates that more man-power needs to be devoted to this subsystem in future competitions. The design of the recovery system worked nearly perfectly during the full-scale test flight, but failed on competition day which goes to show how unreliable recovery system can be. No one person is at fault for the failure, but to avoid future issues more team members should be primarily devoted to its development. 7.0 Summary of overall experience Overall, our team had a great, exciting, and fulfilling experience participating in the competition. There were many nights and days well spent on writing reports and manufacturing our rocket. Watching our rocket launch was just a stream of emotions that cannot be described. It s great to see something fly that you ve worked for so long on. We were all were pretty frustrated tje main parachute did not deploy during the competition launch, wishing we could have mitigated it beforehand. However, this failure will not stop us from moving forward with future rocket projects and spreading our knowledge down to other students.
15 7.1 Educational Engagement Summary In this section, the student outreach done thus far will be discussed. The final count for the student outreach was 314 students, 310 being Educational/Direct interactions. Below in Table specifics for each event can be seen. The engagement with International Polytechnic High School involved discussing different parameters of the rocket and how they affect performance. These parameters included basic concepts of center of pressure, center of gravity, stability margin and fundamental equations related to performance. In addition, 3-D models of fins and nosecones were brought in and passed around the class. For Almondale and Country Springs Elementary a presentation about the subscale launch vehicle was done. An overall discussion of the launch vehicle components was done. Next the events that go on during a launch were discussed, specifics of the subscale launch for this project were included. For the last portion of the presentation the students came up and assembled the subscale rockets. This included the fins, main parachute bay, FMP bay, avionics bay, drogue parachute and nosecone. Table Educational Engagement Summary Date Location of Event Activity # of students November 28th, 2016 International Polytechnic High School Presentation 4 January 20th, 2017 Almondale Elementary Assembly of subscale rockets 260 February 6th, 2017 Country Springs Elementary Assembly of subscale rockets 25 February 10th, 2017 Country Springs Elementary Assembly of subscale rockets 25 Total 314/ Budget Summary Table shows the overall budget of the Cal Poly Pomona Rocketry team. Most of our costs were allocated toward traveling to the competition. Table shows our funding sources Table Overall Expenses Summary Overall Expenses Cost Launch Vehicle Structure Budget $4, Subscale Launch Vehicle Budget $1, Payload Experiment Budget $1, Other Budget $ Travel Budget $9, TOTAL Full Scale Launch Vehicle cost $5, Total Expenses $16, Table 7.2- Overall Funding Summary Funding Source Date Received Support Accepted Cal Poly Pomona Associated Students Incorporated (ASI) Grant 1/27/2017 $5,500 X California Space Grant 2/3/2017 $11,000 X Scholars Research and Project Grants 3/5/2017 $1,000 X Outside Sources (companies, donations, online funding, etc.) 2/19/2017 $2,000 X Overall Income $19,500 Leftover Income $2,805